Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

CLARK X AIRFOIL (clarkx-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: CLARK X AIRFOIL (clarkx-il)
Reynolds number: 200,000
Max Cl/Cd: 72.35 at α=5.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-clarkx-il-200000-n5.txt
Download as CSV file: xf-clarkx-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: CLARK X AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -11.250  -0.4584   0.10017   0.09628  -0.0413   1.0000   0.0328
 -11.000  -0.6655   0.05211   0.04824  -0.0748   1.0000   0.0335
 -10.750  -0.7068   0.04773   0.04374  -0.0729   1.0000   0.0334
 -10.500  -0.7437   0.04323   0.03893  -0.0696   1.0000   0.0335
 -10.250  -0.7511   0.04017   0.03564  -0.0680   0.9987   0.0340
 -10.000  -0.7216   0.03971   0.03521  -0.0696   0.9956   0.0348
  -9.750  -0.6943   0.03871   0.03412  -0.0715   0.9918   0.0358
  -9.500  -0.6716   0.03642   0.03157  -0.0736   0.9870   0.0370
  -9.250  -0.6500   0.03344   0.02817  -0.0757   0.9825   0.0385
  -9.000  -0.6287   0.03064   0.02484  -0.0770   0.9768   0.0400
  -8.750  -0.6018   0.02833   0.02206  -0.0787   0.9734   0.0412
  -8.500  -0.5768   0.02688   0.02049  -0.0793   0.9682   0.0421
  -8.250  -0.5469   0.02595   0.01947  -0.0806   0.9642   0.0432
  -8.000  -0.5149   0.02504   0.01842  -0.0822   0.9614   0.0445
  -7.750  -0.4879   0.02398   0.01715  -0.0826   0.9561   0.0458
  -7.500  -0.4598   0.02280   0.01574  -0.0832   0.9511   0.0470
  -7.250  -0.4286   0.02168   0.01438  -0.0844   0.9476   0.0482
  -7.000  -0.4007   0.02079   0.01328  -0.0847   0.9422   0.0493
  -6.750  -0.3731   0.01985   0.01220  -0.0849   0.9364   0.0507
  -6.500  -0.3421   0.01908   0.01140  -0.0859   0.9324   0.0522
  -6.250  -0.3147   0.01848   0.01074  -0.0860   0.9262   0.0537
  -6.000  -0.2865   0.01786   0.01004  -0.0862   0.9201   0.0552
  -5.750  -0.2556   0.01723   0.00931  -0.0868   0.9157   0.0570
  -5.500  -0.2298   0.01672   0.00870  -0.0864   0.9080   0.0588
  -5.250  -0.2006   0.01624   0.00811  -0.0867   0.9022   0.0605
  -5.000  -0.1738   0.01562   0.00751  -0.0866   0.8955   0.0633
  -4.750  -0.1464   0.01523   0.00710  -0.0865   0.8883   0.0663
  -4.500  -0.1179   0.01485   0.00665  -0.0866   0.8823   0.0696
  -4.250  -0.0913   0.01452   0.00623  -0.0862   0.8742   0.0724
  -4.000  -0.0633   0.01405   0.00577  -0.0863   0.8682   0.0764
  -3.750  -0.0372   0.01377   0.00547  -0.0859   0.8596   0.0812
  -3.500  -0.0087   0.01349   0.00512  -0.0859   0.8534   0.0870
  -3.250   0.0172   0.01320   0.00487  -0.0855   0.8445   0.0941
  -3.000   0.0454   0.01294   0.00458  -0.0855   0.8379   0.1026
  -2.750   0.0718   0.01272   0.00438  -0.0852   0.8291   0.1147
  -2.500   0.0994   0.01246   0.00415  -0.0850   0.8220   0.1315
  -2.250   0.1258   0.01222   0.00397  -0.0848   0.8131   0.1530
  -2.000   0.1527   0.01193   0.00379  -0.0846   0.8056   0.1865
  -1.750   0.1788   0.01165   0.00372  -0.0843   0.7967   0.2429
  -1.500   0.2057   0.01146   0.00362  -0.0841   0.7886   0.2893
  -1.250   0.2323   0.01127   0.00353  -0.0838   0.7798   0.3298
  -1.000   0.2585   0.01107   0.00346  -0.0834   0.7708   0.3739
  -0.750   0.2843   0.01080   0.00338  -0.0830   0.7620   0.4380
  -0.500   0.3083   0.01046   0.00337  -0.0821   0.7518   0.5328
  -0.250   0.3298   0.00999   0.00339  -0.0805   0.7423   0.6775
   0.000   0.3538   0.00965   0.00351  -0.0787   0.7324   0.8398
   0.250   0.3991   0.00962   0.00354  -0.0817   0.7221   0.9396
   0.500   0.4408   0.00966   0.00349  -0.0844   0.7123   0.9743
   0.750   0.4859   0.00969   0.00344  -0.0880   0.7014   0.9966
   1.000   0.5148   0.00976   0.00343  -0.0883   0.6905   1.0000
   1.250   0.5393   0.00984   0.00342  -0.0876   0.6799   1.0000
   1.500   0.5638   0.00993   0.00343  -0.0868   0.6687   1.0000
   1.750   0.5880   0.01003   0.00347  -0.0861   0.6568   1.0000
   2.000   0.6123   0.01013   0.00351  -0.0853   0.6447   1.0000
   2.250   0.6366   0.01025   0.00356  -0.0845   0.6320   1.0000
   2.500   0.6606   0.01038   0.00361  -0.0836   0.6169   1.0000
   2.750   0.6843   0.01052   0.00367  -0.0827   0.5997   1.0000
   3.000   0.7079   0.01068   0.00375  -0.0818   0.5823   1.0000
   3.250   0.7318   0.01085   0.00385  -0.0809   0.5663   1.0000
   3.500   0.7559   0.01103   0.00398  -0.0801   0.5512   1.0000
   3.750   0.7799   0.01122   0.00412  -0.0794   0.5363   1.0000
   4.000   0.8037   0.01143   0.00427  -0.0785   0.5203   1.0000
   4.250   0.8272   0.01166   0.00444  -0.0777   0.5035   1.0000
   4.500   0.8506   0.01190   0.00464  -0.0768   0.4863   1.0000
   4.750   0.8740   0.01215   0.00485  -0.0760   0.4697   1.0000
   5.000   0.8972   0.01243   0.00507  -0.0751   0.4532   1.0000
   5.250   0.9201   0.01272   0.00533  -0.0742   0.4363   1.0000
   5.500   0.9427   0.01303   0.00560  -0.0732   0.4188   1.0000
   5.750   0.9651   0.01336   0.00588  -0.0723   0.4007   1.0000
   6.000   0.9873   0.01371   0.00620  -0.0713   0.3834   1.0000
   6.250   1.0093   0.01407   0.00654  -0.0703   0.3675   1.0000
   6.500   1.0312   0.01444   0.00690  -0.0693   0.3524   1.0000
   6.750   1.0525   0.01485   0.00728  -0.0683   0.3363   1.0000
   7.000   1.0725   0.01532   0.00769  -0.0670   0.3167   1.0000
   7.250   1.0912   0.01585   0.00815  -0.0656   0.2937   1.0000
   7.500   1.1091   0.01642   0.00861  -0.0641   0.2687   1.0000
   7.750   1.1272   0.01699   0.00910  -0.0626   0.2460   1.0000
   8.000   1.1442   0.01761   0.00963  -0.0610   0.2239   1.0000
   8.250   1.1612   0.01823   0.01020  -0.0595   0.2032   1.0000
   8.500   1.1774   0.01888   0.01079  -0.0578   0.1835   1.0000
   8.750   1.1915   0.01959   0.01143  -0.0559   0.1646   1.0000
   9.000   1.2052   0.02033   0.01211  -0.0538   0.1423   1.0000
   9.250   1.2143   0.02139   0.01298  -0.0514   0.1057   1.0000
   9.500   1.2178   0.02289   0.01420  -0.0483   0.0694   1.0000
   9.750   1.2230   0.02433   0.01551  -0.0456   0.0474   1.0000
  10.000   1.2310   0.02561   0.01677  -0.0434   0.0386   1.0000
  10.250   1.2384   0.02695   0.01811  -0.0412   0.0337   1.0000
  10.500   1.2484   0.02811   0.01939  -0.0394   0.0308   1.0000
  10.750   1.2558   0.02951   0.02086  -0.0375   0.0285   1.0000
  11.000   1.2611   0.03111   0.02254  -0.0355   0.0267   1.0000
  11.250   1.2688   0.03255   0.02412  -0.0340   0.0252   1.0000
  11.500   1.2750   0.03417   0.02586  -0.0325   0.0237   1.0000
  11.750   1.2793   0.03601   0.02779  -0.0310   0.0226   1.0000
  12.000   1.2799   0.03824   0.03011  -0.0296   0.0217   1.0000
  12.250   1.2810   0.04053   0.03250  -0.0284   0.0211   1.0000
  12.500   1.2835   0.04276   0.03486  -0.0274   0.0204   1.0000
  12.750   1.2850   0.04517   0.03739  -0.0265   0.0198   1.0000
  13.000   1.2859   0.04770   0.04005  -0.0258   0.0192   1.0000
  13.250   1.2861   0.05037   0.04282  -0.0253   0.0186   1.0000
  13.500   1.2860   0.05317   0.04573  -0.0249   0.0180   1.0000
  13.750   1.2847   0.05617   0.04883  -0.0247   0.0175   1.0000
  14.000   1.2818   0.05945   0.05219  -0.0246   0.0171   1.0000
  14.250   1.2766   0.06306   0.05587  -0.0247   0.0166   1.0000
  14.500   1.2762   0.06623   0.05917  -0.0248   0.0163   1.0000
  14.750   1.2751   0.06952   0.06259  -0.0250   0.0160   1.0000
  15.000   1.2737   0.07292   0.06612  -0.0252   0.0157   1.0000
  15.250   1.2721   0.07640   0.06973  -0.0256   0.0154   1.0000
  15.500   1.2703   0.07997   0.07342  -0.0261   0.0151   1.0000
  15.750   1.2683   0.08360   0.07717  -0.0267   0.0148   1.0000
  16.000   1.2661   0.08733   0.08103  -0.0275   0.0146   1.0000
  16.250   1.2636   0.09120   0.08502  -0.0284   0.0143   1.0000
  16.500   1.2608   0.09520   0.08914  -0.0295   0.0140   1.0000
  16.750   1.2576   0.09929   0.09333  -0.0307   0.0138   1.0000
  17.000   1.2542   0.10347   0.09762  -0.0322   0.0135   1.0000
  17.250   1.2507   0.10771   0.10196  -0.0337   0.0133   1.0000
  17.500   1.2472   0.11197   0.10629  -0.0353   0.0131   1.0000
  17.750   1.2439   0.11616   0.11056  -0.0368   0.0129   1.0000
  18.000   1.2385   0.12091   0.11544  -0.0388   0.0127   1.0000
  18.250   1.2314   0.12631   0.12102  -0.0414   0.0125   1.0000
  18.500   1.2235   0.13195   0.12684  -0.0443   0.0124   1.0000
  18.750   1.2151   0.13787   0.13294  -0.0476   0.0123   1.0000
  19.000   1.2057   0.14415   0.13940  -0.0511   0.0122   1.0000
  19.250   1.1951   0.15094   0.14638  -0.0552   0.0121   1.0000
<< Back to CLARK X AIRFOIL (clarkx-il)

Polar data table (+)

Polar graphs


<< Back to CLARK X AIRFOIL (clarkx-il)